Known turbine engines, including gas/combustion turbine engines and steam turbine engines, incorporate shaft-mounted turbine blades circumferentially circumscribed by a turbine casing or housing. The remainder of this description focuses on applications within combustion or gas turbine technical application and environment, though exemplary embodiments described herein are applicable to steam turbine engines. In a gas/combustion turbine engine, hot combustion gasses flow in a combustion path that initiates within a combustor and are directed through a generally tubular transition into a turbine section. A forward or Row 1 vane directs the combustion gasses past successive alternating rows of turbine blades and vanes. Hot combustion gas striking the turbine blades cause blade rotation, thereby converting thermal energy within the hot gasses to mechanical work, which is available for powering rotating machinery, such as an electrical generator.
Engine internal components within the hot combustion gas path are exposed to combustion temperatures approximately well over 1000 degrees Celsius (1832 degrees Fahrenheit). The engine internal components within the combustion path, such as for example combustion section transitions, vanes and blades are often constructed of high temperature resistant superalloys. Blades, vanes, and transitions often include cooling passages terminating in cooling holes on component outer surface, for passage of coolant fluid into the combustion path.
Turbine engine internal components often incorporate a thermal barrier coat or coating (“TBC”) of metal-ceramic material that is applied directly to the external surface of the component substrate surface or over an intermediate metallic bond coat (“BC”) that was previously applied to the substrate surface. The TBC provides a thermal insulating layer over the component substrate, which reduces the substrate temperature. Combination of TBC application along with cooling passages in the component further lowers the substrate temperature.
Fabrication of cooling passages in, and application of TBC layers to superalloy components, creates conflicting manufacturing constraints. Traditionally, cooling passages are formed by removing superalloy material from the intended passage path within the component, with exemplary removal tools including mechanical cutting/drilling bits, or various ablation devices, such as high-pressure water jet, percussion laser pulsation, and electric discharge machining (“EDM”). Cut cooling passage path, profile and size are limited by the physical capabilities of the cutting instrument. For example, drilled passages are linear and have cross sectional symmetry to match the drill bit. Ablated passages are limited by the size of the ablation instrument and ability to maneuver the instrument along a cutting path.
Investment cast turbine engine components are fabricated by creating a hardened wax pattern, in a wax injection mold, which replicates the profile of the finished superalloy component. The wax pattern is enveloped in ceramic slurry, which is subsequently hardened by firing, into a ceramic shell casing. When wax is removed from the ceramic shell casing, the internal cavity is filled with molten superalloy material. Typically, more particularly, wax patterns for investment cast, superalloy components for combustion turbine engines, are injected into hard tool wax molds, and removed from the tools with precise and smooth surfaces. The wax patterns are then dipped in various ceramic slurry mixtures and processed to form the ceramic outer shell, which is subsequently sintered to form a vestibule in which molten metal is poured. Upon cooling and solidification, the outer ceramic shell is removed by mechanical and/or chemical methods and the metal part is then prepared for further processing. Further processing of the metal part includes ceramic core removal, finish machining, drilling of cooling holes, and application of a thermal barrier coating (“TBC”). Current state of the art processes often require the investment cast surface be lightly grit blasted to prepare the surface for bond coat application. At this point a bond coat, typically a metallic Cramoium, Aluminum, Yitria (“MCrAlY”) coating is applied to the substrate via a spray deposition technique, such as High Velocity Oxy Fuel (“HVOF”) or Low Pressure Plasma Spray (“LPPS”). After this a ceramic thermal barrier such as YSZ (Yttria Stabilized Zirconia) is applied to the surface of the MCrAlY via atmospheric or air plasma spray (“APS”) to complete the coating system. In some cases, a two layer ceramic coating is applied via APS for low thermal conductivity.
The investment casting wax pattern does not have sufficient, reliable, structural integrity to form cooling passages directly therein. When cooling passages are formed in the mold that forms the wax pattern, there is more than insignificant chance that the cooling passage profile in the wax pattern will deform, or that the wax pattern passage will not fill completely with ceramic slurry; in either case the resultant passage in the metal casting does not confirm to design specification.
In some investment casting, component manufacturing processes, refractory metal core (“RMC”) inserts that conform to the desired profiles and paths of cooling passages are placed in the molds prior to component metal casting. The RMC inserts have to be aligned precisely within the molds, and are removed after casting by chemical dissolution processes, adding to manufacturing complexity and expense. TBC layer application adds additional sequencing challenges to the manufacturing process.
If cooling passages are formed in the blade, vane, transition, or other superalloy component prior to application of the TBC layer, the passages will become obstructed by the TBC material as the latter is applied to the component surface. Obstruction can be mitigated by temporarily masking the cooling passages on the component surface prior to the TBC application, which adds additional, costly, steps to the manufacturing processes. In the alternative, excess TBC material obstructions within cooling passages can be removed subsequently by the aforementioned cutting processes. Post TBC-application cooling passage obstruction removal increases risk of TBC layer damage and/or delamination along the margins of cooling passages on the component surface. In some manufacturing processes, cooling passages are formed after application of a TBC layer to the component substrate. In one known post TBC-coating cooling passage formation process, a pulsed laser ablates TBC material from the component at the intended cooling passage entry point, and then ablates the superalloy material to form the passage.
As previously noted, there is risk of damage to the previously applied TBC layer, or delamination of the layer from the component substrate, as cooling passages are subsequently created within the component. Due to differences in thermal expansion, fracture toughness and elastic modulus, among other things, between typical metal-ceramic TBC materials and typical superalloy materials used to manufacture the aforementioned exemplary turbine components, there is potential risk of thermally- and/or mechanically-induced stress cracking of the TBC layer as well as TBC/turbine component adhesion loss at the interface of the dissimilar materials as the TBC layer and superalloy material are removed during cooling passage formation or cooling passage cleaning to remove TBC obstructions. The cracks and/or adhesion loss/delamination negatively affect the TBC layer's structural integrity and potentially lead to its spallation (i.e., separation of the TBC insulative material from the turbine component).